EP3467257A1 - Vane platform cooling structuresplatform - Google Patents
Vane platform cooling structuresplatform Download PDFInfo
- Publication number
- EP3467257A1 EP3467257A1 EP18186880.3A EP18186880A EP3467257A1 EP 3467257 A1 EP3467257 A1 EP 3467257A1 EP 18186880 A EP18186880 A EP 18186880A EP 3467257 A1 EP3467257 A1 EP 3467257A1
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- EP
- European Patent Office
- Prior art keywords
- combustor
- vane platform
- vane
- aperture
- cavity
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Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/141—Shape, i.e. outer, aerodynamic form
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
- F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/002—Wall structures
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/10—Manufacture by removing material
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/20—Manufacture essentially without removing material
- F05D2230/21—Manufacture essentially without removing material by casting
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/35—Combustors or associated equipment
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
- F05D2240/81—Cooled platforms
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Architecture (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
- This disclosure was made with government support under Contract No. FA8650-15-D-2502/0002 awarded by the United States Air Force. The government has certain rights in the disclosure.
- The present disclosure relates to cooling structures for gas turbine engines and, more specifically, to vane platform cooling structures.
- A gas turbine engine typically includes a fan section, a compressor section, a combustor section, and a turbine section. The fan section may drive air along a bypass flowpath while the compressor section may drive air along a core flowpath. In general, during operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases flow through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads. The interface between a downstream edge of the combustor and an upstream edge of the first vane stage of the high pressure turbine may exhibit premature oxidization.
- A vane platform cooling system is disclosed herein, in accordance with various embodiments. The vane platform cooling system may comprise a vane platform and an orifice formed in a forward end of the vane platform. A first aperture may be formed in a surface of the vane platform. A first channel may be formed through the vane platform and may connect the orifice and the first aperture.
- In various embodiments, the first channel may be configured such that air exits the first aperture at a first angle of less than 20° relative to the surface of the vane platform. A second angle formed by a surface defining the first channel and a plane parallel to the surface of the vane platform may be between 5° and 15°. In various embodiments, a cross-sectional area of the first aperture may be greater than a cross-sectional area of the first channel.
- In various embodiments, a second aperture may be formed in the surface of the vane platform. A second channel may be formed through the vane platform and may connect the orifice and the second aperture. In various embodiments, an airflow structure may be located in the first channel. The airflow structure may comprise at least one of a teardrop shape, an airfoil shape, a pedestal shape, or a racetrack shape.
- In various embodiments, a combustor shell may be located forward of the vane platform. A combustor panel may be coupled to the combustor shell. A cavity may be located between the combustor shell and the combustor panel. A surface defining the cavity may be angled toward the combustor shell. In various embodiments, a ramp may be coupled to an exterior surface of the combustor panel. The ramp may include the surface defining the cavity. A standoff may be located in the cavity. The standoff may comprise at least one of a teardrop shape, an airfoil shape, or a racetrack shape.
- A gas turbine engine is disclosed herein, in accordance with various embodiments. The gas turbine engine may comprise a combustor. The combustor may comprise a combustor shell defining a combustion chamber of the combustor. A combustor panel may be disposed inside the combustion chamber. A cavity may be located between the combustor shell and the combustor panel. A surface defining the cavity may be angled toward the combustor shell. A vane platform may be located aft of an outlet of the combustor. An orifice may be formed in a forward end of the vane platform. A first aperture may be formed in a surface of the vane platform. A first channel may be formed through the vane platform and may connect the orifice and the first aperture.
- In various embodiments, a ramp may be coupled to the combustor panel. The ramp may comprise the surface defining the cavity. In various embodiments, a standoff may be located in the cavity. The standoff may be integral to the ramp. In various embodiments, a plurality of airflow structures maybe located in the first channel. A first airflow structure of the plurality of airflow structures may be configured to direct airflow in a first direction. A second airflow structure of the plurality of airflow structures may be configured to direct airflow in a second direction different from the first direction.
- In various embodiments, a second aperture may be formed in the surface of the vane platform. A second channel may be formed through the vane platform and may connect the orifice and the second aperture.
- A method of cooling a first vane stage platform system of a high pressure turbine is disclosed herein, in accordance with various embodiments. The method may comprise forming an impingement hole through a combustor shell. The impingement hole may allow cooling air to flow into a cavity located between the combustor shell and a combustor panel. The method may further comprise angling a surface of the cavity away from a gap defined by an aft end the combustor panel and a vane platform of the first vane stage platform system, forming an orifice in a forward end of the vane platform, forming an aperture in a surface of the vane platform, and forming a channel fluidly connecting the orifice and the aperture.
- In various embodiments, the method may further comprise forming a plurality of airflow structures in the channel. Each airflow structure of the plurality of airflow structures may comprise at least one of a teardrop shape, an airfoil shape, a pedestal shape, or a racetrack shape. The orifice, the aperture, the channel, and the plurality of airflow structures may be formed during a manufacturing of the vane platform. In various embodiments, angling the surface of the cavity may comprise manufacturing the combustor panel to include a sloped exterior surface. In various embodiments, angling the surface of the cavity may comprise coupling a discrete ramp to an exterior surface of the combustor panel. In various embodiments, the method may further comprise forming the orifice, the aperture, and the channel by at least one of integral casting, electrical discharge machining, laser drilling, or punching through the vane platform.
- The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements.
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FIG. 1 illustrates a cross-sectional view of an exemplary gas turbine engine, in accordance with various embodiments; -
FIG. 2 illustrates a cross-sectional view of an exemplary combustor, in accordance with various embodiments; -
FIG. 3A illustrates a perspective view of a combustor outlet and vane platform interface, in accordance with various embodiments; -
FIG. 3B illustrates a cross-sectional view of the combustor outlet/vane platform interface ofFIG. 3A , in accordance with various embodiments; -
FIG. 3C illustrates a radially inward looking view of the combustor outlet/vane platform interface ofFIG. 3A , in accordance with various embodiments; -
FIG. 4A illustrates a perspective view a combustor outlet and vane platform interface, in accordance with various embodiments; -
FIG. 4B illustrates a cross-sectional view of the combustor outlet/vane platform interface ofFIG. 4A , in accordance with various embodiments; and -
FIG. 5 illustrates a method of cooling a first vane stage platform system of a high pressure turbine, in accordance with various embodiments. - The detailed description of various embodiments herein makes reference to the accompanying drawings, which show various embodiments by way of illustration. While these various embodiments are described in sufficient detail to enable those skilled in the art to practice the disclosure, it should be understood that other embodiments may be realized and that logical aerodynamic, thermodynamic, and mechanical changes may be made without departing from the scope of the disclosure. Thus, the detailed description herein is presented for purposes of illustration only and not of limitation. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented. Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected, or the like may include permanent, removable, temporary, partial, full, and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact.
- Cross hatching lines may be used throughout the figures to denote different parts but not necessarily to denote the same or different materials. Throughout the present disclosure, like reference numbers denote like elements. Accordingly, elements with like element numbering may be shown in the figures, but may not necessarily be repeated herein for the sake of clarity.
- As used herein, "aft" refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine engine. As used herein, "forward" refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion. As used herein, "proximate" refers to a direction inwards, or generally, towards the reference component.
- A first component that is "radially outward" of a second component means that the first component is positioned at a greater distance away from the engine central longitudinal axis than the second component. A first component that is "radially inward" of a second component means that the first component is positioned closer to the engine central longitudinal axis than the second component. In the case of components that rotate circumferentially about the engine central longitudinal axis, a first component that is radially inward of a second component rotates through a circumferentially shorter path than the second component. The terminology "radially outward" and "radially inward" may also be used relative to references other than the engine central longitudinal axis.
- In various embodiments and with reference to
FIG. 1 , agas turbine engine 20 is provided.Gas turbine engine 20 may be a two-spool turbofan that generally incorporates afan section 22, acompressor section 24, acombustor section 26, and aturbine section 28. Alternative engines may include, for example, an augmentor section among other systems or features. In operation,fan section 22 can drive fluid (e.g., air) along a bypass flow-path B whilecompressor section 24 can drive air along a core flow-path C for compression and communication intocombustor section 26 then expansion throughturbine section 28. Although depicted as a turbofangas turbine engine 20 herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures, as well as industrial gas turbines. -
Gas turbine engine 20 may generally comprise alow speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine central longitudinal axis A-A' relative to an enginestatic structure 36 viaseveral bearing systems 38, 38-1, and 38-2. Engine central longitudinal axis A-A' is oriented in the z direction on the provided x-y-z axes. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally be provided, including for example, bearingsystem 38, bearing system 38-1, and bearing system 38-2. -
Low speed spool 30 may generally comprise aninner shaft 40 that interconnects afan 42, alow pressure compressor 44 and alow pressure turbine 46.Inner shaft 40 may be connected to fan 42 through a gearedarchitecture 48 that can drivefan 42 at a lower speed thanlow speed spool 30.Geared architecture 48 may comprise agear assembly 60 enclosed within agear housing 62.Gear assembly 60 couplesinner shaft 40 to a rotating fan structure.High speed spool 32 may comprise anouter shaft 50 that interconnects ahigh pressure compressor 52 and highpressure turbine section 54. Acombustor 56 may be located betweenhigh pressure compressor 52 andhigh pressure turbine 54. In various embodiments, enginestatic structure 36 may include amid-turbine frame 57. Themid-turbine frame 57, if included, may be located generally betweenhigh pressure turbine 54 andlow pressure turbine 46.Mid-turbine frame 57 may support one ormore bearing systems 38 inturbine section 28.Inner shaft 40 andouter shaft 50 may be concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A-A', which is collinear with their longitudinal axes. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine. - The core airflow C may be compressed by
low pressure compressor 44 thenhigh pressure compressor 52, mixed and burned with fuel incombustor 56, then expanded through thehigh pressure turbine 54 andlow pressure turbine 46.Turbines low speed spool 30 andhigh speed spool 32 in response to the expansion. -
Gas turbine engine 20 may be, for example, a high-bypass ratio geared aircraft engine. In various embodiments, the bypass ratio ofgas turbine engine 20 may be greater than about six (6). In various embodiments, the bypass ratio ofgas turbine engine 20 may be greater than ten (10). In various embodiments, gearedarchitecture 48 may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system.Geared architecture 48 may have a gear reduction ratio of greater than about 2.3 andlow pressure turbine 46 may have a pressure ratio that is greater than about five (5). In various embodiments, the bypass ratio ofgas turbine engine 20 is greater than about ten (10:1). In various embodiments, the diameter offan 42 may be significantly larger than that of thelow pressure compressor 44, and thelow pressure turbine 46 may have a pressure ratio that is greater than about five (5:1). Thelow pressure turbine 46 pressure ratio may be measured prior to inlet oflow pressure turbine 46 as related to the pressure at the outlet oflow pressure turbine 46 prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other turbine engines including direct drive turbofans. A gas turbine engine may comprise an industrial gas turbine (IGT) or a geared engine, such as a geared turbofan, or non-geared engine, such as a turbofan, or may comprise any gas turbine engine as desired. -
Low pressure compressor 44,high pressure compressor 52,low pressure turbine 46, andhigh pressure turbine 54 may comprise one or more stages or sets of rotating blades and one or more stages or sets of stationary (i.e., non-rotating) vanes axially interspersed with the rotating blade stages. -
FIG. 2 shows an exemplary cross-section ofcombustor 56 positioned betweenhigh pressure compressor 52 andhigh pressure turbine 54 of agas turbine engine 20.Combustor 56 includes acombustor chamber 102 defined by anouter shell 104 and aninner shell 184.Inner shell 184 may be radially inward ofouter shell 104. The combustorouter shell 104 and the combustorinner shell 184 may provide structural support to thecombustor 56 and its components. For example, a combustorouter shell 104 and a combustorinner shell 184 may comprise a substantially cylindrical or a substantially conical canister portion defining an inner area comprising thecombustor chamber 102. - It may be desirable to protect combustor
outer shell 104 and combustorinner shell 184 from the high temperatures flames and/or combustion gases withincombustion chamber 102. Accordingly, one or more combustor panels 110 (e.g., thermal shields, combustor liners) may be disposed inside thecombustor chamber 102 and may provide such protection. Thecombustor panels 110 may be mounted and/or coupled to thecombustor shell 104/184 via one or more attachment features 106. Thepanels 110 may be spaced apart from the interior surface of their associated shells. For example, one or more outboard combustor panels (e.g.,panels 110a) may be arranged radially inward of the combustorouter shell 104 and one or more inboard combustor panels (e.g.,panels 110b) may also be arranged radially outward of the combustorinner shell 184. -
Combustor panels 110 may be made of any suitable heat tolerant material. In this manner, thecombustor panels 110 may be substantially resistant to thermal mechanical fatigue in order to inhibit cracking of thecombustor panels 110 and/or to inhibit liberation of portions of thecombustor panels 110. In various embodiments, thecombustor panels 110 may be made from a nickel based alloy and/or a cobalt based alloy, among others. For example, thecombustor panels 110 may be made from a high performance nickel-based super alloy. In various embodiments, thecombustor panels 110 may be made from a cobalt-nickel-chromium-tungsten alloy. Thecombustor panels 110 may comprise a partial cylindrical or conical surface section. Thecombustor panels 110 may comprise a variety of materials, such as metal, metal alloys, and/or ceramic matrix composites, among others. - A
diffuser chamber 101 is external thecombustor 56 and cooling air may be configured to flow through thediffuser chamber 101 around thecombustor 56.Combustor chamber 102 may form a region for mixing of core airflow C (with brief reference toFIG. 1 ) and fuel, and may direct the high-speed exhaust gases produced by the ignition of this mixture inside thecombustor 56. The high-speed exhaust gases may be driven downstream within thecombustor 56 to acombustor outlet 160.Combustor outlet 160 may be located immediately ahead (i.e., immediately forward) of a fixedfirst vane stage 162 ofhigh pressure turbine 54. -
First vane stage 162 comprises a plurality ofvane airfoils 120 that are connected by a vaneinner platform 122 and a vaneouter platform 124.First vane stage 162 tends to be the hottest of the vane stages withinhigh pressure turbine 54, asfirst vane stage 162 is closest tocombustor outlet 160. In this regard, vaneinner platform 122, vaneouter platform 124, andairfoils 120 tend to experience an increased and/or earlier occurrence of oxidation and/or other forms of heat damage, as compared to other more downstream (i.e., aft) vane stages. Accordingly, it may be desirable to coolfirst vane stage 162. - With reference to
FIG. 2 andFIGs. 3A ,3B , and3C , features of a vaneplatform cooling system 99 at aninboard interface 100 betweencombustor outlet 160 andfirst vane stage 162 are illustrated, in accordance with various embodiments. WhileFIGs. 3A, 3B , and3C illustrate theinboard interface 100 betweencombustor outlet 160 andfirst vane stage 162, it should be understood that anoutboard interface 103, with momentary reference toFIG. 2 , ofcombustor outlet 160 and first vane stage 162 (i.e., the interface betweenouter platform 124 andouter shell 104 andpanels 110a) may include the elements and functionalities as described herein with respect to coolingsystem 99 at inboard interface 100 (i.e., at the interface betweeninner platform 122 andinner shell 184 andpanels 110b). - In various embodiments, an
annular cooling cavity 117 is formed and/or defined between the combustorinner shell 184 andpanel 110b. As mentioned above, cooling air in thediffuser chamber 101 may enter theannular cooling cavity 117 via impingement holes orapertures 105 formed ininner shell 184. That is, impingement holes 105 may extend from adiffuser side 141 of theinner shell 184 to acombustor side 142 of theinner shell 184 and may supply cooling air 116 (FIG. 3B ) intoannular cooling cavity 117. Upon enteringcavity 117, coolingair 116 may flow downstream (i.e., aft) toward aforward end 132 ofinner platform 122. - In accordance with various embodiments, a plurality of
orifices 127 may be formed inforward end 132 ofinner platform 122. Upon exitingannular cooling cavity 117, aportion 116b of cooling airflow may flow intoorifices 127.Orifices 127 may be fluidly coupled to one or more openings orapertures 118 formed through asurface 123 ofinner platform 122 that is oriented toward core flowpath C. Orifices 127 may be fluidly coupled toapertures 118 via one or more flow paths orchannels 119. - With reference to
FIG. 3B ,channels 119 may include and may be defined, at least partially, by opposingsurfaces portion 116b of coolingair 116 may flow intoorifices 127, throughchannels 119, outapertures 118, and alongsurface 123 ofinner platform 122.Orifices 127,channels 119, andapertures 118 may be configured to create a protective coating or "blanket" of air film oversurface 123 ofinner platform 122, thereby protectinginner platform 122 from the hot combustion gases exitingcombustor outlet 160. For example, channels 119 (e.g., surfaces 150 and/or surfaces 152) may be configured such thatportion 116b of coolingair 116 exitsaperture 118 at an angle theta (θ) of 30° or less, relative to surface 123 ofplatform 122. In various embodiments, angle θ may be 20° or less. In various embodiments, angle θ may be 15° or less. Stated differently, in various embodiments, an angle alpha (α) formed bysurface 152 and aplane 156 parallel to surface 123 may be between 1° and 30°, 5° and 20°, or 10° and 17°. The focusedair exiting apertures 118 may flow alongsurface 123 ofinner platform 122 and may create a film over theinner platform 122 andindividual vane airfoils 120 offirst vane stage 162. - In various embodiments, more than one
channel 119 and/or more than oneaperture 118 may be associated (i.e., in fluid communication with) asingle orifice 127. For example, with reference toFIG. 3C ,channel 119a,aperture 118a,channel 119b, andaperture 118b are each in fluid communication withorifice 127a. - With continued reference to
FIG. 3C , in various embodiments, one ormore pins 180 andairflow structures 190 may be formed alongchannels 119.Pins 180 andairflow structures 190 may extend fromsurface 150 to surface 152, with momentary reference toFIG. 3B .Airflow structures 190 may directportion 116b of coolingair 116 in a desired direction.Airflow structures 190 may comprise a tear drop shape (as shown), an airfoil shape (similar to the shape of airfoil 120), a race track or oval shape, or any other desired geometry. In this regard, the number, size, location, shape, and orientation ofairflow structures 190 may be selected based on a desired cooling profile forvane stage 162. For example, the shape and orientation ofairflow structures 190 may be selected todirect portion 116b of coolingair 116 toward areas ofvane stage 162 where increased cooling airflow is desired (i.e., areas subject to increased temperatures or thermal stress) and away from areas where less cooling air flow is needed. For example,airflow structures 190 may directportion 116b of coolingair 116 towardairfoils 120 and areas ofsurface 123 proximate toairfoils 120. Stated another way,airflow structures 190 may be used to increase a volume or flow rate of coolingair 116 in a first direction and decrease the volume or flow rate of coolingair 116 in a second direction. -
Pins 180 may provide structure support and/or an increased surface area for heat exchange and improved cooling. Pins may also directportion 116b of coolingair 116 in a desired direction. In this regard, the number, size, and location ofpins 180 may be selected based on a desired cooling profile forvane stage 162. For example, pins 180 of larger diameter (e.g., pins 180a) or a great density of pins 180 (i.e., a greater number ofpins 180 in a particular area) may be located where a decreased cooling airflow is desired, and pins 180 having a smaller diameter (e.g., pins 180b) or a decreased density ofpins 180 may be located in areas where increased cooling airflow is desired. For example, pins 180 may be used to direct coolingair 116 over areas ofsurface 123 that are subject to increased temperatures or thermal stresses. Stated another way, pins 180 may be used to impedeportion 116b of coolingair 116 from flowing in certain directions, thereby increasing the volume or flow rate ofportion 116b of coolingair 116 in other directions.Pins 180 comprise a generally cylindrical or pedestal shape, though a square geometry, rectangular geometry, elliptical geometry, or other shaped geometry may also be employed. - In various embodiments,
orifices 127,channels 119,apertures 118, pins 180, andairflow structures 190 may be formed during a manufacturing ofinner platform 122. In this regard,inner platform 122 may be formed using, for example, a machining, integral casting, molding, or additive manufacturing process, andorifices 127,channels 119,apertures 118, pins 180, andairflow structure 190 may be formed during said process. - In various embodiments, and with reference to
FIGs. 3A and 3B , an opening orgap 130 may be defined between anaft end 115 ofcombustor panel 110b andforward end 132 ofinner platform 122.Gap 130 may be an axial space betweencombustor panel 110b and inner platform 122 (e.g.,combustor panel 110b is axially spaced apart from inner platform 122).Gap 130 may fluidly connectannular cooling cavity 117 to core flow path C. Said differently, aportion 116a of coolingair 116 may flow fromcavity 117, through thegap 130, and mix with the combustiongases exiting outlet 160. - It may be desirable to minimize the amount or
portion 116a of coolingair 116 escaping throughgap 130, and thereby maximize the amount orportion 116b of coolingair 116 supplied toorifices 127 and exitingapertures 118. In this regard, proximate toaft end 115, anexterior surface 114 ofcombustor panel 110b may be sloped or angled towardinner shell 184. As used herein, an "exterior" surface of a combustor panel refers to a combustor panel surface that is oriented away fromcombustion chamber 102, with momentary reference toFIG. 2 . The slope ofexterior surface 114 tends to direct the coolingair 116 exiting the aft end ofinterior cooling cavity 117 away fromgap 130. The slope or angle ofexterior surface 114 towardsurface 142 ofinner shell 184 may reduce theportion 116a of coolingair 116 flowing through channeling 130, thereby increasing theportion 116b of coolingair 116 flowing intoorifice 127. - In various embodiments, a plurality of panel standoffs 198 may be located in
cavity 117.Panel standoffs 198 may be integrally formed withexterior surface 114 ofcombustor panel 110b. In various embodiments, panel standoffs 198 may be formed during a manufacturing ofcombustor panel 110b. In this regard,combustor panel 110b may be formed using, for example, a machining, casting, molding, or additive manufacturing process, and panel standoffs 198 may be formed during said process.Panel standoffs 198 may extend fromexterior surface 114 towardinner shell 184.Panel standoffs 198 may be located as close tosurface 142 ofinner shell 184 as possible. Stated differently, a radial distance betweenpanel standoffs 198 andsurface 142 ofinner shell 184 tends to be minimized to decrease a volume of cooling air flow flowing betweenpanel standoffs 198 andsurface 142. -
Panel standoffs 198 may be located downstream (i.e., aft) ofholes 105.Panel standoffs 198 may be configured to direct coolingair 116 towardorifices 127 andforward end 132 ofinner platform 122.Panel standoffs 198 may comprise a tear drop shape (e.g., 198a), a airfoil shape (e.g. 198b), a race track or oval shape (e.g., 198c), or any other desired geometry. The number, size, location, shape, and orientation of panel standoffs 198 may be selected based on a desired cooling profile forvane stage 162. For example, the shape and orientation of panel standoffs 198 may be selected to concentrate coolingair 116 in orifices 127 (and thus in apertures 118) configured to cool the areas ofvane stage 162 that are subject to increased temperatures and/or thermal stress, anddirect cooling air 116 away fromorifices 127 andapertures 118 that are associated with areas ofvane stage 162 in need of less cooling. - In various embodiments, a
seal 212 may be located betweeninner shell 184 andinner platform 122.Seal 212 may prevent coolingair 116 from avoidingorifices 127 by blockingcooling air 116 from flowing radially inward betweeninner shell 184 andinner platform 122. -
FIGs. 4A and 4B illustrate a vaneplatform cooling system 299 at an inboard interface between acombustor outlet 260 and afirst vane stage 262, in accordance with various embodiments.Combustor outlet 260 may be located at an aft end of acombustor panel 210, similar tocombustor outlet 160 andcombustor panels 110 inFIG. 2 .Combustor panel 210 may be coupled to aninner combustor shell 284.First vane stage 262 comprises a plurality ofvane airfoils 220 which are connected by a vaneinner platform 222 and an outer vane platform, similar to vaneouter platform 124 inFIG. 2 . WhileFIGs. 4A and 4B illustrate the inboard interface ofcombustor outlet 260 andfirst vane stage 262, it should be understood that an outboard interface betweencombustor outlet 260 and first vane stage 262 (i.e., an interface between the outer vane platform and an outer combustor shell and panel) may include the elements and functionalities as described herein with respect to coolingsystem 299 at the inboard interface (i.e., at the interface between vaneinner platform 222 andinner combustor shell 284 and combustor panel 210). - An
annular cooling cavity 217 is formed and/or defined between combustorinner shell 284 andcombustor panel 210.Cooling air 216 may entercavity 217 via impingement holes orapertures 205 formed in theinner combustor shell 284. Upon enteringcavity 217, coolingair 216 may flow downstream (i.e., aft) toward aforward end 232 ofinner platform 222. - In accordance with various embodiments, a plurality of
orifices 227 may be formed inforward end 232 ofinner platform 222. Upon exiting annular cooling cavity 217 aportion 216b of coolingair 216 may flow intoorifices 227.Orifices 227 may be fluidly coupled to one ormore apertures 218.Apertures 218 may be formed through asurface 223 ofinner platform 222.Surface 223 may be oriented toward core flowpath C. Orifices 227 may be fluidly coupled toapertures 218 via one or more flow paths orchannels 219. A diameter oforifices 227, as measured circumferentially, may be greater than a diameter orchannels 219. In this regard, multiple channels 219 (and thus multiple apertures 218) may be in fluid communication with asingle orifice 227. -
Channels 219 andapertures 218 may be configured such thatportion 216b of coolingair 216 exitsaperture 218 at an angle θ, relative tosurface 223, between 1° and 30°. In various embodiments, angle θ is between 5° and 20°. In various embodiments, angle θ is between 10° and 17°. Stated differently, in various embodiments, an angle α formed bychannels 219 and a plane parallel to the z-axis may be between 1° and 30°, 5° and 20°, or 10° and 17°. In various embodiments, a cross-sectional area of eachaperture 218 may be greater than a cross-sectional area of eachchannel 219. The increased cross-sectional area ofapertures 218, as compared to the cross-sectional area ofchannels 219, tends to allow theportion 216b of coolingair 216 exitingapertures 218 to spread acrosssurface 223. The increased cross-sectional area ofapertures 218, as compared to the cross-sectional area ofchannels 219, also tends decrease a flow rate (i.e., a volume of fluid passing per unit time, e.g., liters per second) of the cooling air as it exitsapertures 218. The increased cross-sectional area ofapertures 218, as compared the cross-sectional area ofchannels 219, may thus increase the surface area of the cooling air film oversurface 223, thereby increasing the area protected from the hot combustion gases exitingcombustor outlet 260. - In various embodiments,
orifices 227,channels 219, andapertures 218 may be formed after a manufacturing ofinner platform 222. In this regard,orifices 227,channels 219, andapertures 218 may be incorporated (i.e., retrofitted) into existinginner platforms 222. For example,orifices 227,channels 219, andapertures 218 may be formed by electrical discharge machining (EDM), laser drilling, punching, or any other suitable manufacturing process. - In various embodiments, a
gap 230 may be defined betweenaft end 215 ofcombustor panel 210 andforward end 232 ofinner platform 222.Gap 230 may be an axial gap. Aportion 216a of coolingair 216 may flow fromcavity 217, throughgap 230, and mix with the combustiongases exiting outlet 260. To minimize theportion 216a of coolingair 216 escaping throughgap 230, aramp 231 may be located incavity 217.Ramp 231 may be coupled to anexterior surface 214 of combustor panel 210 (i.e., to a surface ofcombustor panel 210 that is oriented towardinner combustor shell 284 and away from the combustion chamber).Ramp 231 may comprise an angle or slopedsurface 233.Surface 233 may be angled or sloped toward asurface 242 ofinner combustor shell 284. The angle or slope ofsurface 233 may direct the coolingair exiting cavity 217 away fromgap 230. Directing coolingair 216 away fromgap 230 may reduce theportion 216a of coolingair 216 escaping throughgap 231, and thereby increase theportion 216b of coolingair 216 supplied toorifices 227 and exitingapertures 218. -
Ramp 231 may be discrete fromcombustor panel 210. Stated differently, ramp 231 may be formed in a manufacturing process separate fromcombustor panel 210. In this regard,ramp 231 may be incorporated (i.e., retrofitted) onto existing combustor panels.Ramp 231 may be coupled to the exterior surface ofcombustor panel 210 in any suitable manner. For example, ramp 231 may be welded or brazed tocombustor panel 210. In various embodiments, fasteners (e.g., clips, screws, bolts, etc.) may employed to coupledramp 231 tocombustor panel 210. - In accordance with various embodiments,
standoffs 234 may be located incavity 217.Standoffs 234 may be integrally formed withramp 231. In various embodiments,standoffs 234 may be coupled (e.g., welded or brazed) to surface 233 oframp 231.Standoffs 234 may extend fromsurface 233 towardinner combustor shell 284.Standoffs 234 are located as close tosurface 242 ofinner combustor shell 284 as possible. Stated differently, a radial distance betweenstandoffs 234 andsurface 242 is minimized to decrease a volume of coolingair 216 flowing betweenstandoffs 234 andsurface 242. -
Standoffs 234 may be located downstream (i.e., aft) ofholes 205 and may direct coolingair 216 towardforward end 232 ofinner platform 222 and towardorifices 227.Standoffs 234 may comprise a tear drop shape, a airfoil shape, a race track or oval shape, or any other desired geometry. The number, size, location, shape, and orientation ofstandoffs 234 may be selected based on a desired cooling profile forvane stage 262. For example, the shape and orientation ofstandoffs 234 may be selected to concentrate coolingair 216 inorifices 227 and/orapertures 218, which create cooling film over the areas ofvane stage 262 that are most susceptible to thermal stress, and/or to direct coolingair 216 away fromorifices 227 andapertures 218 that are associated with cooler areas ofvane stage 262.Standoffs 234 may introduce swirl or may direct coolingair 216 straight toward theforward end 232 ofinner panel 222.Standoffs 234, in combination with the size, location, and number ofholes 205, may meter the flow rate of coolingair 216. -
Ramp 231 andstandoffs 234 may increase the amount of cooling air supplied toorifices 227 andapertures 218.Orifices 227,channels 219, andapertures 218 may cause theportion 216b of coolingair 216 exitingapertures 218 to form a cooling film alongsurface 233.Orifices 227,channels 219, andapertures 218 may also direct theportion 216b of coolingair 216 exitingapertures 218 toward areas that experience increased oxidation or other thermal stresses. Controlling a cooling of the vane platform may result in increased part life, which tends to decrease costs. - Referring to
FIG. 5 , amethod 300 of cooling a first vane stage of a high pressure turbine is illustrated, in accordance with various embodiments.Method 300 may comprise forming an impingement hole through a combustor shell (step 302). The impingement hole may allow cooling air to flow into a cavity located between the combustor shell and a combustor panel.Method 300 may further comprise angling a surface of the cavity away from a gap defined by an aft end the combustor panel and a vane platform of the first vane stage (step 304). Astep 306 ofmethod 300 may comprise forming an orifice in a forward end of the vane platform, forming a aperture in a surface of the vane platform, and forming a channel fluidly connecting the orifice and the aperture. - In various embodiments, step 306 of
method 300 may further comprise forming a plurality of airflow structures in the channel. Each airflow structure of the plurality of airflow structures may comprise at least one of a teardrop shape, an airfoil shape, or a racetrack shape. The orifice, the aperture, the channel, and the plurality of airflow structures may be formed during a manufacturing of the vane platform. In various embodiments,step 304 may comprise manufacturing the combustor panel to include a sloped exterior surface. In various embodiments,step 304 may comprise coupling a discrete ramp to an exterior surface of the combustor panel. In various embodiments,step 306 may further comprise forming the orifice, the aperture, and the channel by at least one of electrical discharge machining, laser drilling, or punching through the vane platform. - Benefits, other advantages, and solutions to problems have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, solutions to problems, and any elements that may cause any benefit, advantage, or solution to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the inventions. The scope of the inventions is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean "one and only one" unless explicitly so stated, but rather "one or more." Moreover, where a phrase similar to "at least one of A, B, or C" is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C. Different cross-hatching is used throughout the figures to denote different parts but not necessarily to denote the same or different materials.
- Systems, methods and apparatus are provided herein. In the detailed description herein, references to "one embodiment", "an embodiment", "various embodiments", etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments.
- Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f) unless the element is expressly recited using the phrase "means for." As used herein, the terms "comprises," "comprising," or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.
Claims (15)
- A vane platform cooling system, comprising:a vane platform;an orifice formed in a forward end of the vane platform;a first aperture formed in a surface of the vane platform; anda first channel formed through the vane platform and connecting the orifice and the first aperture.
- The vane platform cooling system of claim 1, wherein the first channel is configured such that air exits the first aperture at a first angle of less than 20° relative to the surface of the vane platform.
- The vane platform cooling system of claim 2, wherein a second angle formed by a surface defining the first channel and a plane parallel to the surface of the vane platform is between 5° and 15°.
- The vane platform cooling system of claim 2 or 3, wherein a cross-sectional area of the first aperture is greater than a cross-sectional area of the first channel.
- The vane platform cooling system of any preceding claim, further comprising:a second aperture formed in the surface of the vane platform; anda second channel formed through the vane platform, wherein the second channel connects the orifice and the second aperture.
- The vane platform cooling system of any preceding claim, further comprising an airflow structure located in the first channel.
- The vane platform cooling system of claim 6, wherein the airflow structure comprises at least one of a teardrop shape, an airfoil shape, a pedestal shape, or a racetrack shape.
- The vane platform cooling system of any preceding claim, further comprising:a combustor shell located forward of the vane platform;a combustor panel coupled to the combustor shell; anda cavity located between the combustor shell and the combustor panel, wherein a surface defining the cavity is angled toward the combustor shell.
- The vane platform cooling system of claim 8, further including a ramp coupled to an exterior surface of the combustor panel, wherein the ramp includes the surface defining the cavity.
- The vane platform cooling system of claim 8 or 9, further comprising a standoff located in the cavity, the standoff comprising at least one of a teardrop shape, an airfoil shape, or a racetrack shape.
- A gas turbine engine, comprising:a combustor comprising,a combustor shell defining a combustion chamber of the combustor,a combustor panel disposed inside the combustion chamber, anda cavity located between the combustor shell and the combustor panel, wherein a surface defining the cavity is angled toward the combustor shell; anda vane platform located aft of an outlet of the combustor;an orifice formed in a forward end of the vane platform;a first aperture formed in a surface of the vane platform; anda first channel formed through the vane platform and connecting the orifice and the first aperture.
- The gas turbine engine of claim 11, further comprising a ramp coupled to the combustor panel, wherein the ramp comprises the surface defining the cavity,
optionally further comprising a standoff located in the cavity, wherein the standoff is integral to the ramp, and/or
optionally further comprising a plurality of airflow structures located in the first channel, wherein a first airflow structure of the plurality of airflow structures is configured to direct airflow in a first direction, and wherein a second airflow structure of the plurality of airflow structures is configured to direct airflow in a second direction different from the first direction, and/or
optionally further comprising:a second aperture formed in the surface of the vane platform; anda second channel formed through the vane platform and connecting the orifice and the second aperture. - A method of cooling a first vane stage platform system of a high pressure turbine, comprising:forming an impingement hole through a combustor shell, wherein the impingement hole allows cooling air to flow into a cavity located between the combustor shell and a combustor panel;angling a surface of the cavity away from a gap defined by an aft end the combustor panel and a vane platform of the first vane stage platform system;forming an orifice in a forward end of the vane platform;forming an aperture in a surface of the vane platform; andforming a channel fluidly connecting the orifice and the aperture.
- The method of claim 13, further comprising forming a plurality of airflow structures in the channel, wherein each airflow structure of the plurality of airflow structures comprises at least one of a teardrop shape, an airfoil shape, a pedestal shape, or a racetrack shape, and wherein the orifice, the aperture, the channel, and the plurality of airflow structures are formed during a manufacturing of the vane platform.
- The method of claim 13 or 14, wherein angling the surface of the cavity comprises manufacturing the combustor panel to include a sloped exterior surface, or
wherein angling the surface of the cavity comprises coupling a discrete ramp to an exterior surface of the combustor panel, and / or
optionally further comprising forming the orifice, the aperture, and the channel by at least one of integral casting, electrical discharge machining, laser drilling, or punching through the vane platform.
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US15/727,816 US11118474B2 (en) | 2017-10-09 | 2017-10-09 | Vane cooling structures |
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EP3800328A1 (en) * | 2019-10-04 | 2021-04-07 | Raytheon Technologies Corporation | Engine turbine support structure |
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CN112855285B (en) * | 2019-11-28 | 2023-03-24 | 中国航发商用航空发动机有限责任公司 | Turbine blade and aircraft engine |
DE102020202089A1 (en) * | 2020-02-19 | 2021-08-19 | Siemens Aktiengesellschaft | Platform structure for a turbine blade and additive manufacturing process |
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EP1138878A2 (en) * | 2000-03-31 | 2001-10-04 | ALSTOM Power N.V. | Flat freestanding gas turbine element |
US20050100437A1 (en) * | 2003-11-10 | 2005-05-12 | General Electric Company | Cooling system for nozzle segment platform edges |
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